Combustion Section Heat Transfer System for a Propulsion System

ABSTRACT

The present disclosure is directed to a propulsion system including a wall defining a combustion chamber inlet, a combustion chamber outlet, and a combustion chamber therebetween, a nozzle assembly disposed at the combustion chamber inlet, the nozzle assembly configured to provide a fuel/oxidizer mixture to the combustion chamber, a turbine nozzle coupled to the wall and positioned at the combustion chamber outlet, wherein the turbine nozzle defines a cooling circuit within the turbine nozzle, and a casing positioned radially adjacent to the wall, wherein a channel structure is positioned between the casing and the wall, the channel structure in fluid communication with the cooling circuit within the turbine nozzle, and wherein a flowpath is formed between the wall and the casing, the flowpath in fluid communication from the cooling circuit at the turbine nozzle to the nozzle assembly to provide a flow of oxidizer to the thereto.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation of U.S. application Ser. No.15/623,773 “Combustion Section Heat Transfer System for a PropulsionSystem” having a filing date of Jun. 15, 2017, and which is incorporatedherein by reference in its entirety.

FIELD

The present subject matter relates generally to a system of heattransfer for a combustion section of a propulsion system.

BACKGROUND

Conventional combustion systems generally include a combustion chamberand turbine nozzle exposed to combustion gases downstream of thecombustion chamber. Conventional propulsion systems and combustionsections generally remove an amount of air from the primary flowpath toprovide cooling to the turbine nozzle to mitigate structuraldeterioration. The extracted air from the primary flowpath generallybypasses the combustion chamber. As such, propulsion systems mustbalance utilizing air from the primary flowpath for mitigatingstructural deterioration of the turbine nozzle with performance andefficiency losses due to removing air from mixture with a fuel forgenerating combustion gases.

Therefore, a combustion system is needed that mitigates structuraldeterioration of the turbine nozzle while minimizing or eliminatingperformance and efficiency losses due to utilizing air from the primaryflowpath for heat transfer.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

The present disclosure is directed to a propulsion system including anannular inner wall and an annular outer wall, a nozzle assembly, aturbine nozzle, and an inner casing and an outer casing. The inner walland outer wall together extend at least partially along a longitudinaldirection and together define a combustion chamber inlet, a combustionchamber outlet, and a combustion chamber therebetween. The nozzleassembly is disposed at the combustion inlet and provides a mixture offuel and oxidizer to the combustion chamber. The turbine nozzle definesa plurality of airfoils in adjacent circumferential arrangement disposedat the combustion chamber outlet. The turbine nozzle is coupled to theouter wall and the inner wall. The inner casing is disposed inward ofthe inner wall and the outer casing is disposed outward of the outerwall. Each of the inner casing and the outer casing are coupled to theturbine nozzle. A primary flowpath is defined between the inner casingand the inner wall, through the turbine nozzle, and between the outercasing and the outer wall, and in fluid communication with thecombustion chamber.

In one embodiment, the outer casing is further coupled to the innerwall.

In another embodiment, the outer wall, the inner wall, the turbinenozzle, the inner casing, and the outer casing are an integralstructure.

In various embodiments, the primary flowpath defines an inner diameter(ID) primary flowpath between the inner wall and the inner casing, andan outer diameter (OD) primary flowpath between the outer wall and theouter casing. In one embodiment, the turbine nozzle defines a firstcooling circuit within the plurality of airfoils. The first coolingcircuit is in fluid communication with the primary flowpath.

In another embodiment, the turbine nozzle includes one or more coolingcircuit walls defining the first cooling circuit.

In various embodiments, the first cooling circuit defines a first inletopening in direct fluid communication with the ID primary flowpath. Thefirst cooling circuit defines a first outlet opening in direct fluidcommunication with the OD primary flowpath.

In still another embodiment, the primary flowpath defines a second IDprimary flowpath between the inner casing and the inner wall. The firstcooling circuit defines a first inlet opening in direct fluidcommunication with the ID primary flowpath, and the first coolingcircuit defines a first outlet opening in direct fluid communicationwith the second ID primary flowpath.

In yet another embodiment, the primary flowpath defines a second ODprimary flowpath between the outer casing and the outer wall. The firstcooling circuit defines a first inlet opening in direct fluidcommunication with the OD primary flowpath, and the first coolingcircuit defines a first outlet opening in direct fluid communicationwith the second OD primary flowpath.

In still various embodiments, the turbine nozzle defines a secondcooling circuit within the plurality of airfoils. The second coolingcircuit is in direct fluid communication with one or more of the IDprimary flowpath and the OD primary flowpath. In one embodiment, thesecond cooling circuit is disposed at a trailing edge of the pluralityof airfoils. In another embodiment, the propulsion system defines a hotgas path downstream of the combustion chamber, in which the secondcooling circuit is in direct fluid communication with the hot gas path.

In various embodiments, a channel structure is defined in one or more ofthe ID primary flowpath and the OD primary flowpath. The channelstructure is in direct fluid communication with the first coolingcircuit. In one embodiment, the channel structure comprises one or morechannel walls extended at least along the longitudinal directiondefining one or more cooling channels in fluid communication with thefirst cooling circuit and the primary flowpath. In another embodiment,the one or more channel walls are coupled to the one or more coolingcircuit walls. The first cooling circuit is in direct fluidcommunication with the cooling channels.

In various embodiments, a heat exchanger is disposed within the primaryflowpath. In one embodiment, the heat exchanger is disposed within oneor more of an ID primary flowpath, a second ID primary flowpath, an ODprimary flowpath, and a second OD primary flowpath.

In another embodiment, a support member is extended at least partiallyalong the radial direction from one or more of the outer casing and theinner casing to one or more of the outer wall and the inner wall. In oneembodiment, the support member defines a passage through which a flow ofoxidizer from an upstream inlet section enters one or more of an innerdiameter (ID) primary flowpath and an outer diameter (OD) primaryflowpath.

In various embodiments, the outer wall, the inner wall, and nozzleassembly together define at least in part a rotating detonationcombustion system.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic view of a propulsion system in accordance with anexemplary embodiment of the present disclosure;

FIG. 2 is a cross sectional view of an exemplary embodiment of acombustion system of the propulsion system of FIG. 1;

FIG. 3 is a cross sectional view of another exemplary embodiment of acombustion system of the propulsion system of FIG. 1;

FIG. 4 is a cross sectional view of yet another exemplary embodiment ofa combustion system of the propulsion system of FIG. 1;

FIG. 5 is a cross sectional view of still another exemplary embodimentof a combustion system of the propulsion system of FIG. 1; and

FIG. 6 is a cross sectional view of still yet another exemplaryembodiment of a combustion system of the propulsion system of FIG. 1;

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

Embodiments of a combustion system are generally provided that maymitigate structural deterioration of a turbine nozzle while minimizingor eliminating performance and efficiency losses due to utilizing anoxidizer from a primary flowpath for heat transfer. Various embodimentsof the combustion system generally provided and described herein utilizeregenerative cooling via a closed-loop arrangement of the oxidizer froman inlet section through a turbine nozzle and to a nozzle assembly andcombustion chamber. In one embodiment, the combustion system defines adetonation combustion system, in which efficiency and performance may befurther improved by decreasing or minimizing detonation cell size in thecombustion chamber via the increased temperature of the oxidizer fromthe turbine nozzle.

Referring now to the figures, FIG. 1 depicts a propulsion system 102including a combustion system 100 in accordance with an exemplaryembodiment of the present disclosure. The propulsion system 102generally includes an inlet section 104 and an outlet section 106, withthe combustion system 100 located downstream of the inlet section 104and upstream of the exhaust section 106. In various embodiments, thepropulsion system 102 defines a gas turbine engine, a ramjet, or otherpropulsion system including a fuel-oxidizer burner producing combustiongases that provide propulsive thrust or mechanical energy output. In anembodiment of the propulsion system 102 defining a gas turbine engine,the inlet section 104 includes a compressor section defining one or morecompressors generating an overall flow of oxidizer 195 to the combustionsystem 100. The inlet section 104 may generally guide a flow of theoxidizer 195 to the combustion system 100. The inlet section 104 mayfurther compress the oxidizer 195 before it enters the combustion system100. The inlet section 104 defining a compressor section may include oneor more alternating stages of rotating compressor airfoils. In otherembodiments, the inlet section 104 may generally define a decreasingcross sectional area from an upstream end to a downstream end proximateto the combustion system 100.

As will be discussed in further detail below, at least a portion of theoverall flow of oxidizer 195 is mixed with a fuel to generate combustiongases 138. The combustion gases 138 flow downstream to the exhaustsection 106. In various embodiments, the exhaust section 106 maygenerally define an increasing cross sectional area from an upstream endproximate to the combustion system 100 to a downstream end of thepropulsion system 102. Expansion of the combustion gases 138 generallyprovides thrust that propels the apparatus to which the propulsionsystem 102 is attached, or provides mechanical energy to one or moreturbines further coupled to a fan section, a generator, or both. Thus,the exhaust section 106 may further define a turbine section of a gasturbine engine including one or more alternating rows or stages ofrotating turbine airfoils. The combustion gases 138 may flow from theexhaust section 106 through, e.g., an exhaust nozzle 135 to generatethrust for the propulsion system 102.

As will be appreciated, in various embodiments of the propulsion system102 defining a gas turbine engine, rotation of the turbine(s) within theexhaust section 106 generated by the combustion gases 138 is transferredthrough one or more shafts or spools 110 to drive the compressor(s)within the inlet section 104. In various embodiments, the inlet section104 may further define a fan section, such as for a turbofan engineconfiguration, such as to propel air across a bypass flowpath outside ofthe combustion system 100 and exhaust section 106.

It will be appreciated that the propulsion system 102 depictedschematically in FIG. 1 is provided by way of example only. In certainexemplary embodiments, the propulsion system 102 may include anysuitable number of compressors within the inlet section 104, anysuitable number of turbines within the exhaust section 106, and furthermay include any number of shafts or spools 110 appropriate formechanically linking the compressor(s), turbine(s), and/or fans.Similarly, in other exemplary embodiments, the propulsion system 102 mayinclude any suitable fan section, with a fan thereof being driven by theexhaust section 106 in any suitable manner. For example, in certainembodiments, the fan may be directly linked to a turbine within theexhaust section 106, or alternatively, may be driven by a turbine withinthe exhaust section 106 across a reduction gearbox. Additionally, thefan may be a variable pitch fan, a fixed pitch fan, a ducted fan (i.e.,the propulsion system 102 may include an outer nacelle surrounding thefan section), an un-ducted fan, or may have any other suitableconfiguration.

Moreover, it should also be appreciated that the combustion system 100may further be incorporated into any other suitable aeronauticalpropulsion system, such as a turboshaft engine, a turboprop engine, aturbojet engine, a ramjet engine, a scramjet engine, etc. Further, incertain embodiments, the combustion system 100 may be incorporated intoa non-aeronautical propulsion system, such as a land-based ormarine-based power generation system. Further still, in certainembodiments, the combustion system 100 may be incorporated into anyother suitable propulsion system, such as a rocket or missile engine.With one or more of the latter embodiments, the propulsion system maynot include a compressor in the inlet section 104 or a turbine in theexhaust section 106.

Referring now to FIGS. 1-6, the combustion system 100 generally includesan annular outer wall 118 and an annular inner wall 120 spaced from oneanother along the radial direction R and extended generally along thelongitudinal direction L. The outer wall 118 and the inner wall 120together define in part a combustion chamber inlet 124, a combustionchamber outlet 126, and a combustion chamber 122 therebetween.

The combustion system 100 further includes a nozzle assembly 128disposed at the combustion chamber inlet 124. The nozzle assembly 128provides a flow mixture of oxidizer 195 and a liquid or gaseous fuel (orcombination thereof) to the combustion chamber 122, wherein such mixture(shown schematically as arrows 132) is combusted/detonated to generatecombustion gases 138 therein. The combustion gases 138 then flow fromthe combustion chamber 122 downstream through the combustion chamberoutlet 126 to the exhaust section 106.

The nozzle assembly 128 includes one or more nozzles 140. In oneembodiment, the nozzle 140 defines an annular nozzle around thelongitudinal centerline 116, in which the flow of oxidizer 195 passesacross a generally converging-diverging nozzle wall of the nozzle 140. Aplurality of fuel injection ports are defined in adjacentcircumferential arrangement along the annulus of the nozzle 140 aroundthe longitudinal centerline 116. The oxidizer 195 passes through thenozzle 140 and mixes with a liquid or gaseous fuel (or combinationthereof).

In another embodiment, the nozzle 140 defines a plurality of nozzles(e.g., fuel injectors) disposed in adjacent arrangement along thecircumferential direction C. The nozzle assembly 128 may include aplurality of discrete nozzles 140, each including a fuel injection portand nozzle flowpath through which the oxidizer 195 and fuel mixes.

In either embodiment, the nozzle assembly 128 may further define aplurality of nozzles 140 arranged in adjacent arrangement along theradial direction R. Each nozzle 140 of the nozzle assembly 128 maydefine a plurality of geometric structures specific to one or moreoperating conditions of the propulsion system (e.g., a power output, apressure or temperature condition at the combustion system 100, etc.),or one or more structures attenuating pressure oscillations, pressurewaves, acoustics, etc. from the combustion/detonation of thefuel/oxidizer mixture 132 in the combustion chamber 122.

Referring still to FIGS. 1-6, the combustion system 100 further includesa turbine nozzle 200 defining a plurality of airfoils 215 in adjacentarrangement along the circumferential direction C. The turbine nozzle200 is disposed at the combustion chamber outlet 126. The turbine nozzle200 is coupled to the outer wall 118 and the inner wall 120 of thecombustion system 100. The turbine nozzle 200, or more specifically, theplurality of airfoils 215 of the turbine nozzle 200, is configured toaccelerate a circumferential flow component of combustion gases 138 fromthe combustion chamber 122 to the exhaust section 106. For example, theexhaust section 106 may define a turbine section including one or moreturbine rotors. The turbine nozzle 200 is configured to accelerate acircumferential flow component of the combustion gases 138 flowinggenerally along the longitudinal direction L such as to reduce a normalforce of the combustion gases 138 on the turbine rotors downstream ofthe turbine nozzle 200.

The combustion system 100 further includes an inner casing 203 disposedinward of the inner wall 120 along the radial direction R and an outercasing 201 disposed outward of the outer wall 118 along the radialdirection R. Each of the inner casing 203 and the outer casing 201extend at least partially along the longitudinal direction L and definea generally annular structure around the longitudinal centerline 116.Each of the inner casing 203 and the outer casing 201 extend toward theturbine nozzle 200 and are each coupled thereto. For example, the innercasing 203 and the outer casing 201 are each coupled toward a downstreamend of the turbine nozzle 200.

The propulsion system 102 defines a primary flowpath 250 through theinlet section 104, the combustion system 100, and the exhaust section106. Within the combustion system 100, the primary flowpath 250 isdefined between the inner casing 203 and inner wall 120, through theturbine nozzle 200, between the outer casing 201 and outer wall 118, andthrough the nozzle assembly 128 and between the outer wall 118 and innerwall 120 (i.e., through the combustion chamber 122).

During operation of the propulsion system 102, a flow of oxidizer 195 isingested through the inlet section 104 and provided to the combustionsystem 100. The oxidizer 195 flows from an upstream end from the inletsection 104 through the primary flowpath 250 within the combustionsystem 100. Referring to FIGS. 1-3, the oxidizer 195, in serial order,flows through the primary flowpath 250 defined between the inner casing203 and the inner wall 120, through the turbine nozzle 200 (e.g.,through one or more passages defined within the plurality of airfoils215 of the turbine nozzle 200), through the primary flowpath 250 betweenthe outer casing 201 and the outer wall 118, through the nozzle assembly128 and combustion chamber 122, then egressing the combustion system 100into the exhaust section 106 across the turbine nozzle 200. As theoxidizer 195 flows through the nozzle assembly 128, it is then mixedwith a fuel and ignited within the combustion chamber 122 to generatecombustion gases 138. As such, the combustion system 100 may minimize oreliminate a portion of oxidizer utilized exclusively for cooling (i.e.,minimize or eliminate a portion of oxidizer removed from mixing withfuel and generation of combustion gases).

The flow of oxidizer 195 through the primary flowpath 250 as describedherein provides convective cooling of the combustion system 100,including the outer wall 118, the inner wall 120, and the turbine nozzle200. The primary flowpath 250 providing convective cooling of thecombustion system 100 may improve combustion system 100 durability andefficiency, and improve turbine nozzle 200 durability.

Referring now to FIG. 2, a cross sectional view of an exemplaryembodiment of the combustion system 100 is generally provided. Invarious embodiments, the outer casing 201 is coupled to the inner wall120. For example, the outer casing 201 or inner casing 120 may include asupport member 205 extended at least partially along the radialdirection R. The support member 205 may define a curve or contouredstructure (e.g., a dome, wall, vane, etc.) configured to direct the flowof oxidizer 195 toward the primary flowpath 250 through the combustionsystem 100, such as between the inner casing 203 and the inner wall 120.

The primary flowpath 250 through the combustion system 100 may furtherdefine an inner diameter (ID) primary flowpath 251 between the innerwall 120 and the inner casing 203 and an outer diameter (OD) primaryflowpath 253 between the outer wall 118 and the outer casing 201.

In one embodiment of the combustion system 100, such as generallyprovided in FIG. 2, the turbine nozzle 200 defines a first coolingcircuit 210 within the plurality of airfoils 215. The first coolingcircuit 210 is defined by one or more cooling circuit walls 212 extendedwithin the plurality of airfoils 215. The one or more cooling circuitwalls 212 defines the first cooling circuit 210 as one or more channelsextended generally at least along a radial direction R within theturbine nozzle 200. However, it should be appreciated that in otherembodiments, cooling circuit walls 212 may define the first coolingcircuit 210 along the longitudinal direction L, in a serpentinestructure, or further at least partially along a circumferentialdirection, etc.

The first cooling circuit 210 is in fluid communication with the primaryflowpath 250. In one embodiment, such as generally provided in FIG. 2,the first cooling circuit 210 defines a first inlet opening 211 indirect fluid communication with the ID primary flowpath 251 and a firstoutlet opening 213 in direct fluid communication with the OD primaryflowpath 253. The first inlet opening 211 adjacent to the ID primaryflowpath 251 enables a flow of oxidizer 195 to enter from the inletsection 104 through the ID primary flowpath 251 into the first coolingcircuit 210 through the plurality of airfoils 215 of the turbine nozzle200. The first outlet opening 213 adjacent to the OD primary flowpath253 enables the flow of oxidizer 195 to egress from the first coolingcircuit 210 through the OD primary flowpath 253 and through the nozzleassembly 128 and into the combustion chamber 122. The flow of oxidizer195 as described herein enables heat transfer to the oxidizer 195 suchas to provide cooling to the inner wall 120, the turbine nozzle 200, andthe outer wall 118.

Referring now to FIG. 3, a cross sectional view of another exemplaryembodiment of the combustion system 100 is generally provided. Thecombustion system 100 shown in FIG. 3 may be configured substantiallysimilarly as described in regard to FIG. 2. However, the embodimentprovided in FIG. 3 further defines a second cooling circuit 220 withinthe plurality of airfoils 215 of the turbine nozzle 200. The secondcooling circuit 220 is defined by one or more of the cooling circuitwalls 212 within the plurality of airfoils 215.

The second cooling circuit 220 defines a second inlet opening 221 indirect fluid communication with the ID primary flowpath 251 and a secondoutlet opening 223 in direct fluid communication with a hot gas path255. The hot gas path 255 is defined as the portion of the primaryflowpath 250 downstream of the nozzle assembly 128, such as at anddownstream of a combustion/detonation of the fuel/oxidizer mixture 132within the combustion chamber 122. The second inlet opening 221 adjacentto the ID primary flowpath 251 enables a flow of oxidizer 195 to enterfrom the inlet section 104 through the ID primary flowpath 251 into thefirst cooling circuit 210 and the second cooling circuit 220 through theplurality of airfoils 215 of the turbine nozzle 200. The second outletopening 223 adjacent to the hot gas path 255 enables the flow ofoxidizer 195 to egress from the second cooling circuit 220.

In various embodiments, the second cooling circuit 220 is disposed at atrailing edge of the plurality of airfoils 215 (e.g., a downstream endof the plurality of airfoils 215).

Referring now to FIG. 4, a cross sectional view of yet another exemplaryembodiment of the combustion system 100 is generally provided. Thecombustion system shown in FIG. 4 may be configured substantiallysimilarly as described in regard to FIGS. 2-3. However, the embodimentprovided in FIG. 4 further defines a channel structure 300 within theprimary flowpath 200 in direct fluid communication with the firstcooling circuit 210 of the turbine nozzle 200. The channel structure 300includes one or more channel walls 312 defining one or more coolingchannels 310 within the channel structure 300. As shown in FIG. 4, thechannel walls 312 may extend at least partially along the longitudinaldirection L. In various embodiments, the channel walls 312 are coupledto the cooling circuit walls 212 of the turbine nozzle 200. As such, theone or more cooling channels 310 defined by the channel structure 300are in direct fluid communication with the one or more passages of thefirst cooling circuit 210.

In one embodiment, such as generally provided in FIG. 4, the channelstructure 300 is disposed in the ID primary flowpath 251 between theinner casing 203 and the inner wall 120. In another embodiment, such asgenerally provided in FIG. 5, the channel structure 300 is disposed inthe OD primary flowpath 253 between the outer casing 201 and the outerwall 118. Furthermore, in reference to FIG. 5, the channel structure 300is disposed in the ID primary flowpath 251 and the OD primary flowpath253, each in direct fluid communication with the first cooling circuit210. Although not depicted, in other embodiments the channel structure300 may be disposed solely in the OD primary flowpath 253.

During operation of the propulsion system 102, the combustion system 100enables heat transfer from the turbine nozzle 200 and the inner wall120, outer wall 118, or both. The generally cooler flow of oxidizer 195from the inlet section 104 through the combustion section 100 enablesheat transfer from the generally hot outer wall 118, inner wall 120, andturbine nozzle 200 to the flow of oxidizer 195. The channel structure300, including the channel walls 312, and the first cooling circuit 210,including the cooling circuit walls 212, enable dissipation of heat fromthe flow of oxidizer 195 as it passes through the cooling channels 310and the first cooling circuit 210.

The embodiments of the combustion system 100 generally provided enableutilization of the flow of oxidizer 195 to cool the turbine nozzle 200,the outer wall 118, and the inner wall 120 while further utilizing theoxidizer 195 for combustion/detonation within the combustion chamber 122while minimizing or eliminating thermodynamic and efficiency losses thatgenerally result from utilizing the oxidizer for cooling rather thancombustion/detonation.

Referring now to FIG. 5, a cross sectional view of another exemplaryembodiment of the combustion system 100 is generally provided. Thecombustion system 100 shown in FIG. 5 may be configured substantiallysimilarly as described in regard to FIGS. 2-4. However, in FIG. 5, aspreviously described, the channel structure 300 is disposed within theID primary flowpath 251 and the OD primary flowpath 253. The combustionsystem 100 defines a second ID primary flowpath 252 between the channelstructure 300 and the inner casing 203. The combustion system 100further defines a second OD primary flowpath 254 between the channelstructure 300 and the outer casing 201.

Referring still to FIG. 5, the first cooling circuit 210 defines thefirst inlet opening 211 in direct fluid communication with the ODprimary flowpath 253 and at least another of the first inlet opening 211in direct fluid communication with the second ID primary flowpath 252.The first cooling circuit 210 further defines the first outlet opening213 in direct fluid communication with the ID primary flowpath 251 andat least another of the first outlet opening 213 in direct fluidcommunication with the second OD primary flowpath 254.

The first inlet opening 211 adjacent to the ID primary flowpath 251enables a flow of oxidizer 195 to enter from the inlet section 104through the ID primary flowpath 251 into the first cooling circuit 210through the plurality of airfoils 215 of the turbine nozzle 200, andthrough the first outlet opening 213 adjacent to the second ID primaryflowpath 252. The first inlet opening 211 adjacent to the OD primaryflowpath 253 enables a flow of oxidizer 195 to enter from the inletsection 104 through the OD primary flowpath 253 into the first coolingcircuit 210 through the plurality of airfoils 215 of the turbine nozzle200, and through the first outlet opening 213 adjacent to the second ODprimary flowpath 254.

The combustion system 100 may further define a support member 205extended from the outer wall 118 to the outer casing 201. In variousembodiments, combustion system 100 defines the support member 205extended from the inner wall 120 to the inner casing 203. The supportmember 205 may define one or more passages 355 through which a flow ofoxidizer 195 from the inlet section 104 enters the ID primary flowpath251, the OD primary flowpath 253, or both. The support member 205further defines a wall extended from the outer casing 201 to the outerwall 118, or from the inner casing 203 to the inner wall 120, or both,in which the wall fluidly segregates the flow of oxidizer 195 from theinlet section 104 from direct fluid communication with the second IDprimary flowpath 252 and the second OD primary flowpath 254. The walldefined by the support member 205 further fluidly segregates the flow ofoxidizer 195 egressing the first cooling circuit 210 to the nozzleassembly 128 from direct fluid communication with flow of oxidizer 195at the primary flowpath 250 upstream of the ID primary flowpath 251 andOD primary flowpath 253.

Referring now to FIG. 6, a cross sectional view of a combustion system100 is generally provided. The combustion system 100 may be configuredsubstantially similarly as described in regard to FIGS. 2-5. However, inFIG. 6, the combustion system 100 further includes a heat exchanger 370disposed within the primary flowpath 250. In the embodiment generallyprovided in FIG. 6, the heat exchanger 370 is disposed within theprimary flowpath 250 between the outer casing 201 and the outer wall118. In other embodiments, the heat exchanger 370 may be disposed withinthe primary flowpath 250 between the inner casing 203 and the inner wall120. In various embodiments, the heat exchanger may be disposed withinone or more of the ID primary flowpath 251, the second ID primaryflowpath 252, the OD primary flowpath 253, and the second OD primaryflowpath 254.

The heat exchanger 370 may include a plurality of fins, plates, walls,tubes or manifolds, or combinations thereof through and to which theoxidizer 195 transfers heat. In one embodiment, such as generallyprovided in FIG. 6, the heat exchanger 370 is disposed within theprimary flowpath 250 downstream of the first cooling circuit 210 (e.g.,at the OD primary flowpath 253, the second OD primary flowpath 254, orthe second ID primary flowpath 252). One or more of the turbine nozzle200, the outer wall 118, and the inner wall 120 transfers thermal energyto the oxidizer 195 as it flows through the primary flowpath 250 fromthe inlet section 104 through the turbine nozzle 200. The oxidizer 195transfer thermal energy to the heat exchanger 370, thereby cooling theoxidizer 195, before entering the nozzle assembly 128 and mixing with aliquid or gaseous fuel for combustion/detonation in the combustionchamber 122.

In other embodiments, the heat exchanger 370 may be disposed within theprimary flowpath 250 upstream of the first cooling circuit 210 (e.g., atthe ID primary flowpath 251 or the OD primary flowpath 253). In stillvarious embodiments, the heat exchanger 370 defines a cooling conduit,including an inlet and an outlet each in direct fluid communication withthe primary flowpath 250, through which the oxidizer 195 flows.

In various embodiments, the heat exchanger 370 includes one or more heattransfer fluids, such as, but not limited to, air, an inert gas, theliquid or gaseous fuel (or combination thereof), oil, lube, or hydraulicfluid, or a liquid or gaseous refrigerant.

Referring now to FIGS. 1-6, various embodiments of the combustion system100 may further define a detonation combustion system, such as arotating detonation combustion system. For a combustion system 100defining a rotating detonation combustion system, the increasedtemperature of the oxidizer 195 from the turbine nozzle 200 to thenozzle assembly 128 improves detonation by improving liquid fuelvaporization when mixed with the oxidizer 195. The combustion system 100further improves detonation by decreasing or minimizing detonation cellsize in the combustion chamber 100 via the increased temperature of theoxidizer 195.

In still various embodiments, the combustion system 100 defines anintegral structure. For example, one or more of the outer wall 118, theinner wall 120, the turbine nozzle 200, the inner casing 203, and theouter casing 201 may together define an integral structure. Thecombustion system 100, or portions thereof, may be manufactured usingone or more additive manufacturing, machining, welding, joining, andbonding processes. The combustion system 100 may be formed of one ormore materials suitable of propulsion system hot sections (e.g.,combustion and turbine sections), including, but not limited to, steel,nickel, aluminum, or alloys of each, or a ceramic matrix composite, orcombinations thereof.

The embodiments of the combustion system 100 generally provided anddescribed herein utilize regenerative cooling via a closed-looparrangement of the oxidizer 195 from the inlet section 104 through theturbine nozzle 200 and to the nozzle assembly 128 and combustion chamber122. The combustion system 100 further improves propulsion system 102efficiency and durability by minimizing the amount of oxidizer 195removed from combustion/detonation and utilized for turbine nozzlecooling 200. Furthermore, the combustion system 100 improves propulsionsystem 102 structural life and decreases costs by minimizingdeterioration of the turbine nozzle 200.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A propulsion system comprising: a wall defining acombustion chamber inlet, a combustion chamber outlet, and a combustionchamber therebetween; a nozzle assembly disposed at the combustionchamber inlet, the nozzle assembly configured to provide a fuel/oxidizermixture to the combustion chamber; a turbine nozzle coupled to the walland positioned at the combustion chamber outlet, wherein the turbinenozzle defines a cooling circuit within the turbine nozzle; a casingpositioned radially adjacent to the wall, wherein a channel structure ispositioned between the casing and the wall, the channel structure influid communication with the cooling circuit within the turbine nozzle,and wherein a flowpath is formed between the wall and the casing, theflowpath in fluid communication from the cooling circuit at the turbinenozzle to the nozzle assembly to provide a flow of oxidizer to thethereto.
 2. The propulsion system of claim 1, wherein the channelstructure is in direct fluid communication with an inlet opening formedat the cooling circuit.
 3. The propulsion system of claim 1, wherein theflowpath is in direct fluid communication with an outlet opening formedat the cooling circuit.
 4. The propulsion system of claim 1, wherein thechannel structure comprises a plurality of channel walls extendedco-directional to the wall defining the combustion chamber.
 5. Thepropulsion system of claim 1, wherein the propulsion system isconfigured to provide a flow of oxidizer to the channel structure thento the cooling circuit within the turbine nozzle, and wherein thepropulsion system is configured to provide the flow of oxidizer from thecooling circuit to the flowpath between the casing and the channelstructure.
 6. The propulsion system of claim 5, wherein the propulsionsystem is configured to provide the flow of oxidizer from the flowpathto the nozzle structure, and wherein the flow of oxidizer is mixed witha flow of fuel.
 7. The propulsion system of claim 1, wherein the wallcomprises an inner wall and an outer wall, and wherein the casingcomprises an inner casing positioned inward of the inner wall and anouter casing positioned outward of the outer wall.
 8. The propulsionsystem of claim 7, wherein the channel structure is positioned betweenthe inner wall and the inner casing, and wherein the flowpath isextended between the inner casing and the inner wall.
 9. The propulsionsystem of claim 7, wherein the channel structure is positioned betweenthe outer wall and the outer casing, and wherein the flowpath isextended between the outer casing and the outer wall.
 10. The propulsionsystem of claim 7, wherein the propulsion system comprises an innerflowpath and an outer flowpath, and wherein the channel structurecomprises a first channel structure positioned between the inner walland the inner casing, and wherein the inner flowpath is extended betweenthe inner casing and the inner wall, and wherein the channel structurecomprises a second channel structure positioned between the outer walland the outer casing, and wherein the outer flowpath is extended betweenthe outer casing and the outer wall.
 11. The propulsion system of claim10, wherein the cooling circuit comprises an inlet opening in directfluid communication with a passage formed by the channel structure, andwherein the cooling circuit comprises an outlet opening in direct fluidcommunication with the flowpath.
 12. The propulsion system of claim 11,wherein the cooling circuit comprises an inner inlet opening positionedin fluid communication with the passage of the channel structurepositioned between the inner casing and the inner wall, and wherein thecooling circuit comprises an inner outlet opening positioned in fluidcommunication with the inner flowpath.
 13. The propulsion system ofclaim 11, wherein the cooling circuit comprises an outer inlet openingpositioned in fluid communication with the passage of the channelstructure positioned between the outer casing and the outer wall, andwherein the cooling circuit comprises an outer outlet opening positionedin fluid communication with the outer flowpath.
 14. The propulsionsystem of claim 11, wherein the cooling circuit comprises an inner inletopening positioned in fluid communication with the passage of thechannel structure positioned between the inner casing and the innerwall, and wherein the cooling circuit comprises an inner outlet openingpositioned in fluid communication with the inner flowpath, and whereinthe cooling circuit comprises an outer inlet opening positioned in fluidcommunication with the passage of the channel structure positionedbetween the outer casing and the outer wall, and wherein the coolingcircuit comprises an outer outlet opening positioned in fluidcommunication with the outer flowpath.
 15. The propulsion system ofclaim 1, the propulsion system comprising a heat exchanger positioned inthermal communication with a flow of oxidizer at the flowpath betweenthe casing and the wall.
 16. The propulsion system of claim 1, thepropulsion system comprising a support member extended from the wall tothe casing.
 17. The propulsion system of claim 16, wherein the supportmember forms a passage through which a flow of oxidizer is provided tothe cooling circuit at the turbine nozzle.
 18. The propulsion system ofclaim 16, wherein the support member comprises a support member wallfluidly segregating a flow of oxidizer to the channel structure from theflow of oxidizer at the flowpath between the channel structure and thecasing.
 19. The propulsion system of claim 16, wherein the supportmember fluidly segregates a flow of oxidizer at the nozzle assembly fromthe flow of oxidizer from an inlet section.
 20. The propulsion system ofclaim 1, wherein the wall and the nozzle assembly together at leastpartially form a rotating detonation combustion system.